Line-of-sight guidance system for missiles



DeC- 26, 1967 G. STCHERBATCHEFF 3,360,214

LINE-OF-SIGHT GUIDANCE SYSTEM FOR MISSILES Filed March 8, 1966 4 Sheets-Sheet 1 Fig. 4

Dec. 26, 1967 G. STCHERBATCHEFF 3,360,214

LINE-OF-SIGHT GUIDANCE SYSTEM FOR MISSILES Filed March 8, 1966 4 Sheets-Sheet 2 .n M m p M 4 p Dec. 26, 1967 G. sTcHr-:RBATCHEFF 3,360,214

LINE-OF-SIGHT GUIDANCE SYSTEM FOR MISSILES Filed March 8, 1966 4 Sheets-Sheet 3 Pfg. 5

Dec. 26, 1967 G, STCHERBATCHEFF 3,360,214

LINOF-SIGHT GUIDANCE SYSTEM FOR MISSILES Filed March 8, 1966 4 Sheets-Sheet 4 Fig.

56 APEV United States Patent Oiice 3,360,214 Patented Dec. 26, 1967 3,360,214 LINE-OF-SIGHT GUIDANCE SYSTEM FOR MISSILES Georges Stcherbatchetf, Paris, France, assigner to Nord- Aviation Socit Nationale de Constructions Aeronautiques, Paris, France Filed Mar. 8, 1966, Ser. No. 532,699 Claims priority, application France, Mar. 16, 1965, 9,361, Patent 1,458,137 6 Claims. (Cl. 244-314) ABSTRACT OF THE DISCLOSURE The invention broadly relates to the guiding of aircrafts. It provides a device for stabilizing the transverse trim of an autorotative missile as guided in alignment with the launching station-target axis, with a View further to increase the accuracy of guidance.

The present invention relates to a system for guiding aircraft and more especially jet-propelled autorotative missiles which spin in flight at a constant speed w about their longitudinal axes of revolution (rolling axes).

The invention is more specifically directed to the stabilization of the transverse trim of an autorotative missile as guided in position with respect to the launching station-target axis or line of sight.

It is known to control the Hight of an autorotative missile by transmitting from the ground to the missile orders which produce action on llight-control members working in two planes at right angles, by determining from the ground any deviations from the line of sight and by correcting the orders which are transmitted to the flightcontrol members as a function of said deviations by means of a control loop.

It is an object of the invention to provide a method and device which enable one to guide an autorotative missile with increased accuracy of guidance and rapidity of response. The method according to the invention comprises the steps of measuring the rate of transverse trim of the missile and of correcting the order which is transmitted to t-he flight-control members according to the result of this measurement.

According to a feature of the invention, the said device includes a control chain consisting of a single measuring device and two flight-control members which supply one pitching-'couple component, said control chain, which is driven in rotation by the missile, being intended to operate alternately in pitching motion and yawing motion and being capable of controlling the perfomance of movements both in direction and in depth.

According to another feature of the invention, the comparison of the order with the measurement of' trim velocity with a view to obtaining a signal adapted for controlling the actuation' of the flight-control members -is effected in a device which will be termed hereinafter a switched network and which carries out at the same time the changing of cor-ordinates which is necessary for the purpose of converting signals delined with respect to fixed references into signals which are defined with respect to missile-'based references, and conversely.

The dilferent features and advantages of the invention will become readily apparent from the description which now follows below,.reference being made to the accompanying drawings, in which:

FIG. 1 is a block diagram of a guidance installation in accordance with the invention;

FIG. 2 represents a rate gyroscope which is designed to measure the transverse trim velocity;

FIG. 3 is a digram of a first embodiment of a switched network, alternative embodiment of which are shown in FIGS. 4 and 9;

FIGS. 5, 6 and 7 represent functions which are generated by the switched network and which constitute an approximation of the function which defines the change of co-ordinates; and

FIG. 8 represents a ctitious, non-switched network which is intended to provide a clearer understanding of the operation of the switched network.

The guidance apparatus which is shown in FIG. l comprises a device 1 which is capable of measuring the alignment error, that is to s-ay the diiferen'ce between the actual position of the center of gravity of the missile `and the position which this latter should theoretically have at each instant, taking into account the direction of the target.

A measuring device of this type is already known per se. For example, such a device comprises on the one hand a sighting telescope which a member of the ground crew directs as accurately as possible onto the target and, on the other hand, a missile-locating system. Said locating system in turn comprises a iiigh-t log consisting of anl infra-red radiation source placed at the rear end of the missile and a goniometer, the optical axis of which is common with that of the sighting telescope and which supplies indications of the deviations between the missile and and said optical axis.

The device 1 is connected to an order computer 2 which, starting from the alignment error, produces a direction order By and a depth order Bz.

In the example which is considered above, the computer 2 will for instance comprise a first device for producing a first signal which is proportional to the deviation between the missile and the optical sighting line or axis and a second device for producing a second signal which is a function of the velocity of relative motion of said optical axis, and means for adding the first signal and second signal.

The order signal thus produced is transmitted to the missile receiver 4 by means of a transmitter 3. On board the missile, a co-ordinate-changing device 5 converts the orders dened with respect to a ground-based system of co-ordinates into orders deiined with respect to a missilebased system of co-ordinates.

The equipment carried by the missile additionally comprises a device 6 which takes a measurement of the instantaneous angular velocity of the missile about an axis at right angles to the rolling axis of the missile (measuring axis) and produces an electric signal which is proportional to said angular velocity.

The signals which are delivered from the devices 5 and 6 are applied simultaneously to a sign detector 7 which produces an electric signal having constant amplitude and either positive or negative polarity depending on the sign of the order to be transmitted to the nightcontrol system. Said signal is applied to a night-control system which operates on the all-or-nothing principle and which is capable of applying to the missile constant pitch control torque in a positive or negative direction with respect to an axis at right angles to its rolling axis.

A flight-control system of this type is advantageously of the type described in French Patent No. 1,099,901 as led on August 9, 1948, in the name of the French Government as represented by the Secretary of State for the Armed Forces (Air Force) in respect of: Device for deflecting a high-velocity gas jet discharged from a nozzle. Accordingly, said flight-control system comprises two blades 8a and 8b which are movable from an inactive position in which they do not intercept the jet which is discharged from the nozzle 9 (as shown looking on the rear end) to an active position in which they intercept the jet, respectively, at two diametrically opposite areas of the right cross-section of said nozzle. The motion of the blades 8a and 8b from the inactive to the active position is controlled by electromagnetic devices 10a, 10b which are actuated by the signal either in one direction or in the other depending on the sign of this latter.

The useful effect of the jet-control members is a component of thrust at right angles to the rolling axis as exerted in a radial plane or so-called plane of action of the control members. By reason of the fact that the said plane of action rotates with respect to a fixed reference at the velocity of rotation w of the missile about its rolling axis, any variation in fiight of the missile can be produced or controlled in this manner.

The alignment guidance system which is represented schematically in FIG. 1 constitutes a feed-back loop which tends to reduce to zero any errors in alignment as a result of suitable action exerted on the flight-control system.

In an alignment guidance system of this type, the present invention is mainly directed to the trim-stabilizing unit or stabilizer which is more especially composed of the devices 5-6 and 7.

In short, the purpose of said stabilizer is to reduce any deviation from the line of sight and to improve the precision of guidance and the speed of response to orders, this being obtained by taking into account at each instant the real angular velocity of the missile which is dependent, not only on the order received, but on a number of different disturbing effects. By virtue of the presence of the stabilizer (5-6-7), the order received no longer acts directly on the flight-control members but is compared with the measured angular velocity, and it is the result of this comparison which produces action on the flight-control system.

The stabilizer (5-6-7) performs the function of feedback loop constituted as follows: the missile has at each instant a certain angular velocity about the measuring axis herein above defined; the device 6 measures said angular velocity and generates a signal which combines with the signal of the order subsequently received, and the detector 7 applies to the flight-control system a signal having a sign such that, in the absence of any order, the action of the flight-control system tends to reduce said angular velocity to zero. In fact, a so-called limit cycle is established within the loop and, when no order is transmitted, maintains at zero the mean value of said angular Vvelocity and therefore of the measuring signal. In the presence of an order, it is the mean value of the sum of the measuring signal and of the order signal which is maintained at zero value.

It should be pointed out that the control system which is provided by the stabilizer must be capable of responding to the sum of the frequency of the order and of the frequency of rotation of the missile about its rolling axis. This sum must therefore be smaller than the frequency of self-oscillation of the feed-back system. The speed of response of the jet-control members makes it possible to obtain a high self-oscillation frequency and consequently a satisfactory response of the complete assembly.

A stabilizing device of this type can be produced at low cost and in a small size, with the result that it will prove particularly suitable for small missiles.

Furthermore, the said device is unaffected by aerodynamic asymmetry of the missile and by zero errors of the measuring device. In fact, any continuous disturbance results in a torque which is Caused to rotate about the rolling axis and which consequently has a zero mean value.

Reference being made to FIG. 2, there will now be described one preferred embodiment of the device for measuring angular velocity. The following description, which is intended to permit a clearer understanding of the invention. is in fact concerned with a device of a general type which is already known and which could be replaced by any other device of like design.

The device described hereunder is constituted by a rate gyroscope which is essentially made up of a rotor 12 which spins about an axis 13 forming one of the central axes of symmetry of a frame 14. Said frame is rotatably mounted on two pivots 15-16 which are fixed to the missile along the other central axis of symmetry of the frame 14 which coincides with the rolling axis.

Any angular velocity of the missile about an axis at right angles to the plane of the frame 14 produces a torque which has a tendency to cause said frame to rotate about the pivots 15-16. This rotation is opposed by a spring 17 and a damping device 18. The torque is measured by the displacement of the frame 14, which produces action on a transducer, not shown in the drawings, and said transducer in turn generates a signal which is proportional to the angular velocity to be measured.

Reference being made to FIG. 3, there will now be described one simple embodiment of the device 5 for changing co-ordinates which essentially forms part of the invention.

The device referred to comprises a gyroscope which essentially consists of a gyro rotor 19 which spins about one of the central axes of symmetry of an inner frame 20. The frame 20 is rotatable about its second central axis of symmetry and is mounted on one of the central axes of symmetry of an outer frame 21. Said outer frame is in turn mounted in two bearings 22 and 23 on a second central axis of symmetry which coincides with the rolling axis of the missile. A switch comprises four conductive sectors 24-25-26-27 which are insulated with respect to each other and an insulating shaft 78 fitted with four conductive rings 28-29-30 and 31. The shaft 78 which is oriented along the rolling axis is integral with the side 32 of the frame 21.

The receiver 4 applies potentials By, Bz, -By and Bz respectively to the four sectors 24 to 27 via the rings 28 to 31.

Two pairs of diametrically opposite brushes respectively designated by the reference numerals 33, 34 on the one hand and 35, 36 on the other hand are connected as follows: the brushes 34 and 36 are connected to ground whilst the brushes 33 and 35 are connected to the sign detector 7 as follows: three resistors 37, 38 and 39 are connected in common at one of their extremities to said sign detector 7 whilst the other extremities of said resistors are respectively connected to the device 6 for measuring angular velocity, to the brush 33 and the brush 35.

The gyroscope maintains the sectors 24 to 27 stationary with respect to the ground, whilst the brushes which are associated with the missile carry out a movement of rotation about the rolling axis at the speed w and consequently sweep the conductive sectors.

The resistors 37-38e39 effect the `summation of the gyrometric signal and a certain function of the signals which are respectively picked-up by the brushes 33 and 35, that is to say, the signals iBy and iBz. The brush pitch is determined so that said function should correspond to a sufficient `approximation of the function which defines the change of co-ordinates from ground to missile.

The above-mentioned function is in fact as follows:

B=By cos tut-I-Bz sin wt FIGS. 5 and 6 show that the functions sin wt and cos wt can be replaced with a sufficient approximation in certain cases by functions which assume during one cycle the necessary values O, +1, 0 and 1. Such a result would be achieved with a single pair of brushes. In fact, the addition of a second pair of brushes and resistors supplying the mean value of their potentials makes it possible to provide a sequence of steps which results in a closer approximation of the function sin wt or cos wt; this sequence of steps is shown in FIG. 7, in which the angles indicated relate to the relative displacements of the brushes and the.

numerals relate to the relative amplitudes of the steps which are respectively, during one cycle: +1/2 (through an angle of 30); +1 (through 60);-i-1/2 (through 30); (through 60); -1/2 (through 30); l (through 60); 1/z (through 30); and 0 (through 60).

It has been possible to demonstrate that the function shown in FIG. 7 does not contain a harmonic component 3, which is a satisfactory result.

In FIG. 4, there is shown a more elaborate embodiment of the trim-stabilizing device.

There are again shown in FIG. 4 the sign-detecting device 7, the jet-intercepting members 8a and 8b and the device 6 for measuring the angular velocity of the missile (rate gyroscope).

The gyroscope which forms part of the co-ordinatechanging device and which is designed and arranged as explained earlier in reference to FIG. 3 is not again shown.

FIG. 4 simply shows in a diagrammatic manner the four conductive sectors 24-25-26-27 and the brushes 33-34-35-36 which have already been described.

The co-ordinate-changing device which is illustrated in FIG. 4 comprises a pair of terminals 40-41 which are respectively connected to the sectors 24 and 26 and between which is applied the order signal By and a pair of 'terminals 42-43 which are respectively connected to the sectors 25 and 27 and between which is applied the order signal Bz derived from the receiver 4.

The brushes 34 and 36 are connected to ground whilst the brushes 33 and 35 are connected to the sign detector 7 through resistors 44 and 45 respectively.

The output of the rate gyroscope 6 is connected to the common point of two resistors 46 and 47 through a junction capacitor 48 decoupled by a resistor 49 which is connected to ground.

The resistor 46 is connected to the common point -between the resistor 44 and the brush 33 whilst the resistor 47 is connected to the common .point between the resistor 45 and the brush 35.

A circuit composed of a resistor 50 which is connected 'in parallel with the assembly consisting of a resistor 51 and a series-connected capacitor 52 connects the terminal 41 to the sector 26.

Similarly, a circuit composed of a resistor 53 which is connected in parallel with the assembly consisting of a resistor 54 and a series-connected capacitor 55 connects the terminal 43 to the sector 25.

` In the device of FIG. 3, the response to the orders B is governed by a law having the form:

wherein 0' is the angular velocity which is measured by the rate .gyroscope In order to increase the speed of response and the precision of the stabilizer, it is endeavoured by means of the alternative form of FIG. 4 to establish a law having the .formz B+K1B"=K20 wherein B is the derivative of the orderwith respect to time.

The above formula expresses the fact that the signal B which is introduced within the feed-back loop formed by the stabilizer contains a term which is proportional to the integral of the input 6' of said loop. f This method of introduction of an integral term in a feed-back system is well known per se. It is merely necessary in order to produce the integral term to introduce aresistor-capacitor circuit or so-called imperfect integrator. The application of this known method to the stabiliz- -ing device in accordance with the invention is in no sense self-evident by reason of the fact that the control input is constituted by a value 0' of the angular velocity ,of the missile as measured by reference to a single shaft which also rotates at the angular velocity w as herein above dened.

In order that the integration of said angular velocity which is thus modulated, as it were, at the high frequency w/2'1r, should perform its intended function in the feedback system, it is obviously necessary that said integration should be effected at an angular velocity which has been previously demodulated, that is to say, after a change of co-ordinates which restores thereto the two components H'y and HZ with respect to two ground-based axes of co-ordinates.

Once the integration of the components @'y and 0'Z has been effected, it is then necessary to remodulate the signals By and Bz which contain the integral term or, in other words, to produce by means of another change of co-ordinates a single signal B which is intended to exert action on the Hight-control members.

The device of FIG. 4 performs with a satisfactory approximation this triple function of demodulation-integration-remodulation in a particularly simple manner.

It has already been shown in reference to FIGS. 3, 5, 6 and 7 that the sector and brush switching device as herein above described carries out with sufficient approximation a change of co-ordinates on the signal which is applied thereto. This change of co-ordinates can be carried out by the switching device both in the direction of integration of a signal and in the direction of reading of said signal.

A switching device of this type is therefore capable of converting the signal 0 into two signals which represent a suflicient approximation of the ideal terms ly=0' cos wt and 12:0' sin wt which define the change of co-ordinates at the time of If the above terms are applied to an imperfect integrator i circuit of the resistor-capacitor type, the terms of wt are evidently eliminated on account of their high fre- Y quency, with the result that the terms which are collected at the output of the integrating circuit only contain the integral of the components of 0 with respect to fixed references, thus achieving the result which was initially contemplated.

Once the integration has been effected, the same switching device effects a change of co-ordinates in the direction of remodulation.

With a view to analyzing the operation of the device of FIG. 4 in greater detail, reference will rst be made to FIG. 8 which represents a fictitious device for the assumed purpose of introducing an integral term into a feed-back system which operates wholly in groundbased references. There is again shown in this figure the rate gyroscope 6, the sign detector 7 and the flight-control members Sa and 8b. The order signal B which is supplied from a low-impedance source (not shown) is applied across terminals 56, 57 and transmitted to the sign detector, on the one hand directly through a resistor 58 and, on the other hand, through a circuit which operates for this input as an imperfect shunt constituted by a capacitor 59 and a resistor 60.

The angular velocity signal 0 is transmitted to the sign detector via the circuit 60-59 which plays the part of an imperfect integrator.

A circuit of this type introduces into the control system, not only the integral of the signal 0 but also, and accessorily, the derivative of the order. The introduction of the derivative of the order makesy it possible tov take i intov account vthe lateral accelerations which are required of the missile and improvesthe response of this latter.

' The fictitious circuit of FIG. 8 must be transposed, as

it were, to meet the condition of a rotating missile having a single plane of action of Hight-control members.v

i Althoughv the operation of the transposedcircuit of mit-the order of -By, the photodiodes V66 and @transmit the order +Bz and the photodiodes 67 vand 69.tran..rnit

FIG. 4 can be fully explained only by calculation,` a

simpiiied*explanationwill nevertheless be ventured upon hereunder, it being understood that this explanation is not intended to constitute a strictdeiinition of principle Fand rthat the value of the invention'is not in any sense dependent thereon.

There are again' shown in FIG.; 4 two imperfect inte- `grating circuits as constituted'by the capacitors S2 and 55 and the resistors 46. vand 47. kAs appears from thev figure, as far as the networks comprisingresistors Sli-51 and ,capacitor 52, and resistors 53-54 and capacitor S5 of the stabilizing device which has just been described .v

arecjon'cernedfwith respect to resistors 46 and 47, the

ot the frequency.

The resistors mand 45 are'associated vwith the resistors: j

46:-47 serve, 'as-has already been explained, to supply the mean value vofthe'potentials of the brushes V33 and 35 v in order toobtain steps of the type shown in FIG. 7.

The resistors 'S0-S1 and 53-54 which are associated with the capacitors52-S5 permit the introduction of the derivative of the orders By,-Bz which 'are' applied in thev form 'of floating potentials-across the terminalsv 40-41 and 42-43'.

'Itv should be pointed vout that the switching system Iwhich is vshown inv PIG. 4 is an original feature and that v 1 other applications thereofv could be contemplated. In order to avoid the use of brushes, itis preferable vin vpractice y tov design thev switching system of F'G. 4 in the form of 1 an electronic device of the optical control type as shown in FIG. 9.

The switching elements employed in a device of this type are photodiodes 61 to 69 which are interconnected via resistors 70 to 73 and capacitors 74 and 75.

There again appear in FIG. 9 the same reference numerals as those given in FIG. 4 in order to designate like elements and iBy and iBz designate respectively the direction and depth orders and the `opposite voltages with respect to ground.

The photodiodes have the property of constituting a break in the circuit at the point at which they are inserted when no light shines thereon and of being conductive in both directions when light shines upon them.

The illumination of the photodiodes is effected by a lamp 76 through slits formed in a disc 77 which is caused by the missile gyroscope to rotate about the rolling axis relatively to the missile. To this end, the said disc is secured to the shaft 7S shown in FIG. 3, instead of the conductive sectors and rings.

It is to be understood that the diagram of FIG. 9 does not illustrate the actual number and arrangement of the slits and the respective positions of the disc and the photodiodes, which will be readily designed by the skilled man.

During one revolution (360), the photodiodes 61 and 63 are thus conductive while the disc rotates from 0 to 120, the photodiodes 62 and 64 are conductive from 180 to 300, the photodiodes 66 and 68 are conductive from 90 to 210, the photodiodes 67 and 69 are conductive from 0 to 30 and from 270 to 360 and the photodiode 65 is conductive from 30 to 90, from 120 to 180, from 210 to 270 and from 300 to 360.

lt can be demonstrated that the switched network of FIG. 9 is equivalent to that of FIG. 4.

In fact, it is apparent that the photodiodes 61 and 63 transmit the order ,--By, the photodiodes 62 and 64 trans- During certain portions of a revolution as rherein abovek deiined, one ofthesys'tems of photodiodes 61 tov 64 and, 66` tov 69is 'conductive and the other *system kof photodiodes is blockedy whereas, during other portions of a revolution, both systems are kconductive at the same time.

In each system, only one pair of photodiodesis conduc-v tive at any given moment while the other pair is blocked.

` The photodiode 65 transmits the step-1 when only one of the two systems is conductive and transmits the step 1/2 when both systems are simultaneously.conductive: the sequence of steps represented in FIG. 7 is thus achieved.

The arrangement of the other elements which form part iswithinr the capacity ofv anyone skilled in the art. In

particular, the sign detector can consist of a high-gainv ampliher followed by an amplitudev limiter which con-v trols a bistable device. v v .f f v v vAs, will be readily understood,v a number of different modifications could'b'e made inthe deviceswhich'have rbeen described in the vforegoing without thereby departing either from the scope or the spirit of the invention. It should be pointed out .in vadditirn'x that the diagrams givenin the accompanying drawings have been v simplied in order that the principle of operation 'ot thej devices hereinabove described-may thus bernardo more:

readily apparent. n r What is claimed is: v f

' vr1. An installation comprising a missile which, in

v isin autorotation about its longitiniinal axis anda remote control station for the remote control vof vthe missile, the

vmissile having ight-control means, operable so as to generate a constant pitch control torque *either in a positive or a negative direction with respect to van axis atright u angles tothe said`r longitudinal axis, receiver means for receiving remote control 'orderv signals `from vthe remote control station, said receiverl means having an output; a co-ordinate'changing device connected to the out-put ofv the receiver means, said co-ordinate changing device having an output; measuring means for generating a gyrometric signal which is a function of the angular velocity of the missile, as measured in a plane at right angles to its longitudinal axis, said measuring means having an output; a sign detector having an input connected to the output respectively of the measuring means and of the coordinate changing device, said sign detector having an output connected to the flight-control means; and the remote control station comprising means for measuring the deviation of the position of the missile in llight with respect to the line of sight, computer means for generating order signals as a function of the said deviation, and means for transmitting the said order signals to the missile.

2. An installation as claimed in claim 1, wherein said co-ordinate changing device comprises a gyroscope having inner and outer frames, said outer frame having an axis of symmetry which coincides With the longitudinal axis of the missile; an insulating shaft attached to said frame along said axis of symmetry; first and second pairs of diametrically opposed conductive sectors mounted on said shaft; means connecting the conductive sectors to the output of the receiver means; first and second pairs of diametrically opposed brushes co-operating with the respective conductive sectors, one brush of each pair being grounded to said frame; and circuit means connecting the input of the sign detector to the output of the measuring means and to the other brushes of said brush pairs.

3. An installation as claimed in claim 2, wherein said circuit means include rst, second and third resistors having each first and second terminals, the second terminals of said resistors being connected to the input of the sign detector, the first terminal of the iirst resistor being connected to the output of the measuring means, and the first terminals of the second and third resistors being respectively connected to the other brushes of said brush pairs. 4. An installation as claimed in claim 2, wherein said circuit means includes a capacitor and first, second, third, fourth and fifth resistors, said fourth and fifth resistors respectively connecting the other brushes of said brush pairs to the input of the sign detector, the first resistor and the capacitor having a junction and respectively connecting the frame ground and the output of the measuring means to said junction, the second and third resistors connecting said junction to the respective other brushes of said brush, said means connecting the conductive sectors `to the output of the receiver means including iirst and second diagonal connections across said rst and second pairs of conductive sectors, and a parallel capacitorresistor integrating circuit in each diagonal connection.

5. An installation as claimed in claim 1, wherein said co-ordinate changing device comprises a gyroscope having inner and outer frames, said outer frame having an axis of symmetry which coincides with the longitudinal axis of the missile; a shaft attached to said frame along said axis of symmetry; a slotted disc mounted on said shaft; a light source located on one side of said disc and, 1ocated on the other side of said disc, a switching network including rst, second, third, fourth, fifth, sixth, seventh, eighth and ninth photodiodes adapted for being illuminated according to a predetermined sequence by the light transmitted from said source across the slotted disc, said switching network further including first and second capacitors and first, second, third, and fourth resistors, said first and second resistors having a junction point, said third and fourth resistors having a further junction point,

said first and third resistors connecting the output of the measuring means to the said respective junction points, said second and fourth resistors connecting the input of the sign detector to the said respective junction points, said fifth photodiode connecting the said junction points together, said receiver means having first, second, third, and fourth inputs, said first and third photodiodes serially connecting the iirst output of said receiver means to said junction point, said second and fourth photodiodes serially connecting the second output of said receiver means to said junction point, said eighth and sixth photodiodes serially connecting the third output of said receiver means to said further junction point, said ninth and seventh photodiodes serially connecting the fourth output of said receiver means to said further junction point, said first capacitor being connected across said rst and second photodiodes and said second capacitor being connected across said eighth and ninth photodiodes.

6. An installation as claimed in claim 1, wherein said co-ordinate changing device includes switchin-g means for generating during the successive cycles of the autorotation of the missile, successive stepped control signals for the missile.

References Cited UNITED STATES PATENTS 2/1958 Hall 244-3.1l

BENJAMIN A. BORCHELT, Primary Examiner.

SAMUEL FEINBERG, Examiner.

T. H. WEBB, Assistant Examiner. 

1. AN INSTALLATION COMPRISING A MISSILE WHICH, IN FLIGHT, IS IN AUTORATION ABOUT ITS LONGITUDINAL AXIS AND A REMOTE CONTROL STATION FOR THE REMOTE CONTROL OF THE MISSILE, THE MISSILE HAVING FLIGHT CONTROL MEANS, OPERABLE SO AS TO GENERATE A CONSTANT PITCH CONTROL TORQUE EITHER IN A POSITIVE OR A NEGATIVE DIRECTION WITH RESPECT TO AN AXIS AT RIGHT ANGLES TO THE SAID LONGITUDINAL AXIS, RECEIVER MEANS FOR RECEIVING REMOTE CONTROL ORDER SIGNALS FROM THE REMOTE CONTROL STATION, SAID RECEIVER MEANS HAVNG AN OUTPUT; A CO-ORDINATE CHANGING DEVICE CONNECTED TO THE OUTPUT OF THE RECEIVER MEANS, SAID CO-ORDINATE CHANGING DEVICE HAVING AN OUTPUT; MEASURING MEANS FOR GENERATING A GYROMETRIC SIGNAL WHICH IS A FUNCTION OF THE ANGULAR VELOCITY 